Various types of braking systems are known. For example, hydraulic, pneumatic and electromechanical braking systems have been developed for different applications. The aerospace industry presents unique operational and safety issues with regard to many braking systems. For example, the need for system redundancy in case of a system or component failure is particularly germane to braking operations of an aircraft.
Brake system architectures for aircraft have been developed previously which meet different redundancy, performance and safety requirements. Such architectures include, for example, redundant digital brake control units (BSCUs) which carry out the brake control and antiskid processing functions. In addition, such architectures include, for example, redundant electromechanical actuator controllers (EMACs) which convert commands from the BSCUs to brake actuator forces. Each EMAC provides electrical power to electromechanical brake actuators included within the brakes for the wheels of the aircraft.
FIG. 1 represents such a brake system architecture which has been developed in the past. The architecture, generally designated in FIG. 1 as braking system 30, includes the aforementioned BSCUs and EMACs which are represented collectively as an electromechanical braking controller 60. The controller 60 receives as its primary inputs i) the brake command signals from pilot brake pedal transducers 46 located in the cockpit of the aircraft, and ii) the outputs of torque and wheel speed sensors 62 included as part of a brake 34 on each wheel 36 of the aircraft.
The braking system 30 receives power from three primary power busses and a secondary power buss included within the aircraft. As is known, an aircraft oftentimes will include multiple power busses. In the exemplary embodiment, the aircraft includes primary power busses PWR1, PWR2 and PWRess. Each power buss preferably is independent of one or more of the other power busses to provide a level of redundancy. For example, the power buss PWR1 consists of an alternating-current (AC) power source AC1 and a commonly generated direct-current (DC) power source DC1. Similarly, the power buss PWR2 consists of an AC power source AC2 and a commonly generated DC power source DC2; and the power buss PWRess consists of an AC power source ACess and commonly generated DC power source DCess.
The power buss PWR1 (i.e., AC1 and DC1) may be derived from power generated by the left wing engine in the aircraft, for example. Similarly, the power buss PWR2 (i.e., AC2 and DC2) may be derived from power generated by the right wing engine. In this manner, if the left wing engine or the right wing engine fails, power is still available to the system 30 via the power buss corresponding to the other engine.
The power buss PWRess (i.e., ACess and DCess) may be derived from power generated by the parallel combination of the left wing engine and the right wing engine. In such manner, power from the power buss PWRess will still be available even if one of the engines fail. In addition, DCess can be powered by a battery in case of total loss of the aircraft engines.
More particularly, the aircraft further includes a DC power buss supplied by a battery on board the aircraft. This power is represented by a DChot power source. The battery may be charged via power from one of the other power sources, or may be charged separately on the ground. The DChot power source is configured to provide battery power to the DCess power buss in the event of loss of AC power.
Various circumstances can arise where power from one or more of the power busses may become unavailable. For example, the left wing engine or the right wing engine could fail causing the PWR1 (AC1/DC1) and PWR2 (AC2/DC2) power sources to go down, respectively. Alternatively, power generating equipment such as a generator, inverter, or other form of power converter could fail on one of the respective power busses resulting in the AC1/DC1, AC2/DC2 and/or ACess/DCess power sources becoming unavailable. In addition, a failure can occur in the cabling providing the power from the respective power sources to the system 30, thus effectively causing the respective power source to no longer be available. For this reason, the routing of the power cables for the different busses preferably occurs along different routes throughout the plane to avoid catastrophic failure on all the power buss cables at the same time.
As mentioned above, such previously developed systems have been shown to satisfy system redundancy, performance and safety requirements associated within an aircraft braking system. Nevertheless, there is a desire to improve the capabilities of such braking systems with respect to other possible failures within an aircraft or other vehicle. For example, there is a strong need in the art for a method for partitioning the power buss(es) within the braking system to reduce the risk of impairing or failing a power buss or supply as a consequence of a system or component failure. Moreover, there is a strong need in the art for a method for further maintaining brake control in an emergency or parking mode despite loss of a power buss and or BSCU, for example.